1. Field of the Invention
The present invention relates to the field of combustion technology, and more particularly to a thermal machine gas turbine.
2. Brief Description of the Related Art
Modern industrial gas turbines (IGT) as a rule are designed with annular combustors. In most cases, smaller IGTs are constructed with so-called “can-annular combustors”. In the case of an IGT with annular combustors, the combustion chamber is delimited by the side walls and also by the inlet and discharge planes of the hot gas. Such a gas turbine is shown in FIGS. 1 and 2. The gas turbine 10 which is shown in the detail in FIGS. 1 and 2 has a turbine casing 11 in which a rotor 12 which rotates around an axis 27 is housed. On the right-hand side, a compressor 17 for compressing combustion air and cooling air is formed on the rotor 12, and on the left-hand side a turbine 13 is arranged. The compressor 17 compresses air which flows into a plenum 14. In the plenum, an annular combustor 15 is arranged concentrically to the axis 27 and, on the inlet side, is closed off by a front plate 19 which is cooled with front plate cooling air 20, and on the discharge side is in communication, via a hot gas passage 25, with the inlet of the turbine 13.
Burners 16, which for example are designed as double-cone burners or EV-burners and inject a fuel-air mixture into the combustor 15, are arranged in a ring in the front plate 19. The hot air flow 26 which is formed during the combustion of the mixture reaches the turbine 13 through the hot gas passage 25 and is expanded in the turbine, performing work. The combustor 15 with the hot gas passage 25 is enclosed on the outside, with a space, by an outer and inner cooling shroud 21 or 31 which, by fastening elements 24, are fastened on the combustor 15, 25 and between themselves and the combustor 15, 25 form an annular outer and inner cooling passage 22 or 32 in each case. In the cooling passages 22, 32, cooling air flows in the opposite direction to the hot gas flow 26 along the walls of the combustor 15, 25 into a combustor dome 18, and from there flows into the burners 16 or, as front plate cooling air 20, flows directly into the combustor 15.
The side walls of the combustor 15, 25 in this case are constructed either as shell elements or as complete shells (outer shell 23, inner shell 33). When using complete shells, the necessity of a parting plane (29 in FIG. 2a) arises for installation reasons, which allows an upper half of the shell 23, 33 (upper half-shell 33a in FIG. 2a) to be detached from the lower half (lower half-shell 33b in FIG. 2a), for example in order to install or to remove the gas-turbine rotor 12. The parting plane 29 correspondingly has two parting plane welded seams 30 (FIG. 2a) which, in the example of the type GT13E2 gas turbine constructed by ALSTOM, are located at the level of the machine axis 27 (3 o'clock and 9 o'clock positions).
As already mentioned, the lower and upper half-shells 33a, 33b must be convectively cooled in each case. In order to promote the cooling, the already mentioned cooling shrouds (co-shirts) 21 and 31 are mounted on the half-shell cold side and deflect ambient air and, on account of the combustor pressure drop or burner pressure drop, guide the ambient air over the half-shells and as a result bring about convective cooling.
The cooling shrouds 21, 31 in this case preferably have the following characteristics and functions:                they seal two plenums or chambers;        they must also seal in relation to each other (requiring installation of a sealing lip or overlap);        they are axially-symmetrically constructed, with exception of the parting plane 29;        during installation of the combustor half-shells they must be guided one inside the other in the parting plane;        the cooling shrouds 31 of the combustor inner shells 33a, b must be guided one inside the other on the parting plane 29 in a “blind” manner (no access for a visual inspection of the connecting plane, on account of being covered by the combustor inner shells);        they are able to have cooling holes (for a specific mass flow of cooling air);        they are able to have cooling holes for a possible impingement cooling (for a specific, locally forced cooling of the half-shells);        they must not absorb large axial or radial forces;        they are as a rule not self-supporting, but are mounted on a supporting component;        they must have a large axial and radial movement clearance, especially during transient operating states;        they must be resistant to temperature (fatigue strength-creep strength);        they must be simply and inexpensively producible; and        they are not permitted to have natural vibrations during operation.        
The inner and outer shells 33 or 23 of a gas turbine such as GT13E2 are thermally and mechanically highly stressed during operation. The strength properties of the material of the shells 23, 33 are greatly dependent upon temperature. In order to keep the material temperature below the maximum permissible material temperature level, the shells 23, 33 are convectively cooled. The profiling and the high thermal load close to the turbine inlet (hot gas passage 25) require above all a constantly high heat transfer in this region, even on the cooling air side. This is achieved by impingement cooling in the case of the outer shell 23. Space and flow conditions, and also sealing against a crossflow, are not provided on the inner shell 33 for such impingement cooling. Therefore, conventional convection cooling is resorted to, in which the intensity of the cooling is increased by reduction of the passage height of the cooling passage 32.
The previously used configuration of the inner cooling shroud 31, having two axial plates, on the one hand is contingent upon spacing tolerances and other irregularities, for example in the flow field upstream of the cooling air inlet into the cooling passage, and on the other hand brings about an undesirable reduction of the mass flow of cooling air in the region of the smaller of the two axial plates.